SpaceX has noted that its Falcon 9 first stage has reached a milestone in achieving a better than 20 to 1 mass ratio:

SPACEX ACHIEVES ORBITAL BULLSEYE WITH INAUGURAL FLIGHT OF FALCON 9 ROCKET.Cape Canaveral, Florida – June 7, 2010"The Merlin engine is one of only two orbit class rocket engines developed inthe United States in the last decade (SpaceX’s Kestrel is the other), and isthe highest efficiency American hydrocarbon engine ever built. The Falcon 9first stage, with a fully fueled to dry weight ratio of over 20, has theworld's best structural efficiency, despite being designed to higher humanrated factors of safety."http://www.spacex.com/press.php?page=20100607

The early versions of the Atlas rocket also reached comparably high mass ratios using both "balloon" tank and common bulkhead design, though the latest version, the Atlas 5 first stage, has a poorer mass ratio in not using either of these methods.As described in SpaceX news releases, the Falcon launchers are able to get their high mass ratios because they use both common bulkheads and lightweight aluminum-lithium alloys, instead of the balloon tanks of the earlier Atlas versions.But then I was startled to see that some early Delta rocket first stages, which were kerosene fueled, also had better than 20 to 1 mass ratios, particularly ones using an extra long first stage tank, known as the Delta Thor ELT:

This is notable because these Delta rocket first stages were able to achieve these high mass ratios without using balloon tanks or common bulkheads. Note that the Atlas 5 first stage remember in not using common bulkheads or balloon tanks results a much poorer first stage mass ratio.We'll show the Delta Thor ELT can become a reusable SSTO with a vertical DC-X landing mode by replacing its RS-27 engine with the higher performance NK-33 engine and adding thermal protection and landing systems. The use of the NK-33 will add only 200 kg to the dry mass even though it has nearly twice the thrust. Interestingly the Delta Thor ELT can be made into a SSTO while keeping the vehicle close to the same size of the original DC-X.The original DC-X created quite a stir when it was first flown because it was produced in such a short period, in less than two years, at relatively low cost, less than $60 million, and most importantly it demonstrated quick turnaround with a small ground crew.The DC-X though was only able to make vertical takeoffs to a few thousand feet altitude and vertical landings using hydrogen fuel. To make an orbital version capable of 10,000 kg payload would require a much larger version at over a billion dollar cost, the DC-Y. Even the 1/2-scale version, the DC-X2, would cost $450 million and would only be suborbital using hydrogen fuel even though this 1/2-scale vehicle was twice the size of the DC-X. It is important then that by switching to hydrocarbon fuel that you can get a fully orbital vehicle of close to the size as the DC-X.The Delta Thor ELT had a gross mass of 84,067 kg and an empty mass of 4,059 kg, for a propellant mass of 80,008 kg. The density of kerolox propellant is about 1,000 kg/m^3, so this corresponds to a propellant volume of about 80 m^3. The DC-X had a conical shape with a base about 4.1 m wide and length about 12 m, for a 3 to 1 ratio of length to base. A propellant tank of volume of that of the Delta 1914 first stage, but conically shaped at the same proportions as the DC-X, gives a base of 4.67 and a length of 14 m.According to this, circular-cross section tanks, such as a cone, can get the same propellant mass to tank mass ratio of cylindrical tanks:

Space Access Update #91 2/7/00.The Last Five Years: NASA Gets Handed The Ball, And Drops It."...part of L-M X-33's weight growth was the "multi-lobed" propellant tanks growing considerably heavier than promised.Neither Rockwell nor McDonnell-Douglas bid these; both used provencircular-section tanks. X-33's graphite-epoxy "multi-lobed" liquidhydrogen tanks have ended up over twice as heavy relative to theweight of propellant carried as the Shuttle's 70's vintage aluminumcircular-section tanks - yet an X-33 tank still split open in testlast fall. Going over to aluminum will make the problem worse; X-33's aluminum multi-lobed liquid oxygen tank is nearly four times asheavy relative to the weight of propellant carried as Shuttle'saluminum circular-section equivalent."http://www.space-access.org/updates/sau91.html

Now we have to mass the thermal protection and landing systems. For thermal protection, we'll assume it'll make a ballistic reentry, base first. The base will only be 4.67 meters wide, giving an area of 17 m^2. Using base first reentry we'll have to cover primarily the base only:

Blue Origin New Shepard."A passenger and cargo spacecraft has considerably less need for cross-range."..."As a result, the craft is much "rounder" than the DC-X, optimized for tankage and structural benefits rather than re-entry aerodynamics. It has not been stated if the vehicle is intended to re-enter base-first or nose first, but the former is most likely for a variety of reasons. For one, it reduces heat shield area, and thus weight, covering only the smaller bottom surface rather than the much larger upper portions. The area around the engines would likely require some sort of heat protection anyway, so by using the base as the heat shield the two can be combined. This re-entry attitude also has the advantage of allowing the spacecraft to descend all the way from orbit to touchdown in a base-first orientation, which would seem to offer some safety benefits as well as reducing aero-loading issues."http://en.wikipedia.org/wiki/Blue_Origin_New_Shepard

We'll use the high temperature resistant but low maintenance metallic shingles developed for the X-33:

These have an areal density of 15 kg/m^2. This will require 255 kg to cover the base only. This plus the 200 kg extra mass for the more powerful NK-33 engine brings the dry mass to 4514 kg. The landing gear for an aerial vehicle is commonly taken as 3% of the landed weight:

More energetic fuels than kerosene are also discussed in Dunn's report. Methylacetene for example with altitude compensation gets an average Isp of 352 s. This will allow about 1,450 kg to orbit:

352*9.8ln(1 + 80,008/(5,115 + 1,450)) = 8,897 m/s.

The cost? The original DC-X cost $60 million. Since this reusable kerosene-fueled version is of similar size it might be estimated to be of approx. the same cost. However, there is this surprising cost for the Delta Thor ELT:

Astronautix though is sometimes inaccurate, but I haven't found any other source estimate for the cost of this stage. Typically the cost of the engine is the largest portion of the cost of a rocket stage, so more than half of the $11.6 million would be for the original RS-27 engine. But this would be for more than the price of the more powerful NK-33 currently at $4 million. The metallic shingle TPS though would also be an additional add on to the cost.Still, it is possible the cost could be in the $10 million to $20 million range. Considering we have a reusable launcher with engines that could get perhaps 10 flights and with possibly a 1,450 kg payload capacity, the price per kilo might be as low as $700/kg, or $350/lbs.

Bob Clark

_________________Nanotechnology now can produce the space elevator and private orbital launchers. It now also makes possible the long desired 'flying cars'. This crowdfunding campaign is to prove it:

Why use SSTO orbit? Why not launch using two stages, the first dropping off and being recovered (vertical landing?) Then the second stage is in effect a DC-X style craft, still using vertical descent, but because of the greatly reduced mass, requiring less fuel, so more cargo.

In effect a smaller DC-X on top of a DC-X.

Judicious use of parachutes on both stages should also reduce fuel mass requirements.

SSTO, even using "off the shelf" currently existing componentry as in RGClark's example, greatly reduces the turn around cost/time of a system. Instead of having to service two (or more) vehicles, you only have to do so on one. If you can pull it off technologically, SSTO is inherently more efficient. You aren't hefting extra sets of engines, tankage, and other mass required in each staging.

SSTO, even using "off the shelf" currently existing componentry as in RGClark's example, greatly reduces the turn around cost/time of a system. Instead of having to service two (or more) vehicles, you only have to do so on one. If you can pull it off technologically, SSTO is inherently more efficient. You aren't hefting extra sets of engines, tankage, and other mass required in each staging.

Surely the whole point of staging is the you are not hauling around extra engines, fuel etc, once you have dropped the empty stages?

With an SSTO you have to get the whole craft up there, including all the weight need to just to get halfway. So you need a lot of oomph all the way up. Which is technically not possible at the moment I believe. The servicing point is right though, but I wonder what the difference is between servicing a single, very complex SSTO craft (complex because it needs to be to achieve SSTO - implicit in the fact no-one has built one yet), and servicing two relatively simple stages, which probably have the same design engines, and are generally pretty similar.

Surely the whole point of staging is the you are not hauling around extra engines, fuel etc, once you have dropped the empty stages?

Its a consequence of a propulsion method with restricted ISP and limited power to weight.

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With an SSTO you have to get the whole craft up there, including all the weight need to just to get halfway. So you need a lot of oomph all the way up. Which is technically not possible at the moment I believe.

Yes, true. SSTO will require that "quantum leap" in propulsion and materials science. Which is really the only way we would be realistically talking about it. Otherwise we are discussing science fiction, or at most absurdly small payloads.

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The servicing point is right though, but I wonder what the difference is between servicing a single, very complex SSTO craft (complex because it needs to be to achieve SSTO - implicit in the fact no-one has built one yet), and servicing two relatively simple stages, which probably have the same design engines, and are generally pretty similar.

More advanced does not necessarily mean higher maintenance costs. Better design practices and electronic self diagnosis, sensors, ie; "smart-systems" can take a lot of the grunt work out of even current Rube Goldberg contraptions. But any reusable vehicle is going to have a certain level of maintenance and checks required. Checking the oil, kicking the tires, x-raying structure for stress cracking, etc. The more individual pieces you have in the system, the more you will have to do, irregardless of whether its some far out "anti-gravity" ship or a DC-3.

SSTO fundamentally makes no sense. The idea is purely "wish driven" as is too much of thinking of space access. An ascent consists of 3 environments: [1] atmosphere, drag etc; [2] high acceleration in vacuum; [3] lower acceleration in vacuum. This defines the 3 stage structure for the smallest possible launchers, as with Microlaunchers.

It can be done with 2 stages for LEO, with stage 2 doing most of the acceleration. The partition of delta v for stage 1, 2 being dependent of stage 1 recovery design etc.

Stages 2 and 3 can be combined, and are combined quite often, are they not? That leaves the atmosphere part, which can be mostly alleviated by airlaunch - a zeroth stage, as it were, such as a carrier aircraft or airship. THat, of course, limits the size of the launch vehicles, which is only partially made up for by the increased Isp and reduced drag. Perhaps 100 tonnes GLOM can be accomodated, suggesting payloads in the 1-2 tonne range, though we could possibly push this higher.

Air launch is not practical because a plane cannot rise high enough for vacuum designed upper stages to operate. So a first atmosphere ascent stage is still needed. The cost of an airplane, operating it and the 2nd set of FAA regs make that much more difficult and expensive than a first stage with enough propellant to rise from the ground.

As mentioned in post (can't see it as I type this) said, a single upper stage can reach LEO. But not to escape. I was assuming 3 stages for Microlaunchers because the normal mission for these is escape.

Also, for optimum trajectory the early phase of the vacuum part after leaving stage 1 is to accelerate at over 1 g--I assume 2 for start. Then after a velocity increase of about 2 km/s for LEO or 3 km/s for escape, a 3rd stage with lower acceleration can take over, supply over 50% of the total velocity for the orbit or escape.

I seem to recall a Selenien Boondocks (?) post regarding airlaunch, saying it allows you to shave off 1km/s of your required delta-V - which, given the rocket equation, translates into significant savings.

Of course, if you use an aerostatic station as your zeroth stage (who said an SSTO had to launch from the ground, rather than merely from a somewhat stationary platform?), you can reach higher altitudes and consequently lower pressures. 100mb, 0mb; they're still an awful lot lower pressures than 1000mb...

I still like the look of a TSTO system based partly on the Airship to Orbit system - a suborbital airship departs from an aerostatic station, carrying an orbital craft which will be launched into orbit.

Terraformer's comment re 1 km/s advantage for airlaunch. Approximately 30% more propellant for first stage makes up for ground launch. That propellant and slightly larger tanks, engine is a lot cheaper than an airplane and the 2nd set of FAA regs for air launch.

Also, air launch confines growth of GLOW. Saturn V proves that size of a ground launch rocket is for all practical purposed, unlimited.

Balloon launch mentioned earlier is even worse than airplane launch as there is no control of balloon position and mass restriction is greater.

At sea: forgot about that option. For Microlaunchers there will be a spar buoy mounted version of launch pad in case a land based launch site cannot be found (existing "spaceports" are likely to charge too much per launch).

Airships: If they have the propulsion to buck any wind, I suppose it's an option, but the weight lift capacity will be very limited. Also, the separate aircraft type will still require the separate additional set of FAA regs.

As far as I know John Powell has not launched any powered airships--only unpowered balloons which go with the wind.

Rockets lifting from sea level is still the only practical means for space access in the present, near future.

I'm not sure if they've launched any airships to nearspace yet, but they've certainly built a couple of large airships - Ascender 90 (built for nearspace under an Air Force contract; they decided to move directly to the next vehicle and so A90 became a low altitude testbed) and Ascender 175 (destroyed in a wind accident before it's first free flight; built under Air Force contract). They've also demonstrated maneuvering capability in nearspace, I recall? The trouble with airships is, of course, the low payloads - but, I'm only really interested in low payload reusable vehicles, instead using current heavy lift to get up Lunar infrastructure; for such a one-off cost, it's not worth investing in anything new, however cheap crew + high value cargo is worth it.

Re. the FAA - they're only a problem in the US. Depending on the range of possible orbits, Antarctica is looking interesting as a location for airlaunch (using an infrastructure based on JPA's). You'd have to register your airships in a country, probably the UK, but given the international airspace above Antarctica...

NASA is in a quandary right now about what to do about their manned flight capability. Congress wants this reinstituted quickly but NASA says they can't do it with the money being provided by Congress. In this report Boeing proposes heavy lift launchers using existing components:

One of the proposals is of a manned launcher with the Orion capsule using a shuttle ET propellant tank and four RS-68 engines. This does not use an upper stage but is not a single-stage-to-orbit vehicle because the final push to orbit is made by the onboard thrusters on the Orion spacecraft. However, it is interesting in this report comparison is made to the S-IVB upper stage on the Apollo rocket. I was reminded of a suggestion of Gary Hudson that the S-IVB would be single-stage-to-orbit with significant payload if it used the high efficiency SSME rather than the J-2 engine:

In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg. This would be just enough to carry the crewed version of the Dragon spacecraft without cargo. Boeing's proposal for a manned capsule the CST-100 might be launchable by this also since it is of comparable size and design to the Dragon:

NASA has shown in their crewed spacecraft versions to want to hearken back to Apollo in their use of capsules. This SSTO idea of Hudson would have the advantage of using a proven Apollo component that is already manrated. The SSME's are also already manrated rather than the RS-68 of the Boeing proposal. Because of its small small size compared to the shuttle ET propellant tank it would also be relatively low cost as well as, only needing one SSME engine. In fact it would even be smaller than the Falcon 9, Delta IV, and Atlas V expendable launchers. Note as well NASA is leaning now to using SSME's or their expendable versions rather than the RS-68 for their shuttle derived manned launchers. Hudson in his article stated the S-IVB was designed by the Douglas Aircraft Company, which merged with McDonnell Aircraft to form McDonnell Douglas. It is notable as well that McDonnell Douglas was also the contractor on the DC-X, legendary for its low development cost, quick turnaround time, and small ground crew. NASA in their shuttle-derived launcher studies have focused on getting a cheaper version of the SSME by making an expendable version. However, the greatest advantage of a SSTO is in being reusable. Then I suggest studies be made on the SSME going the opposite direction: how can it be made to be reusable at much reduced maintenance cost? Now the SSME's have to be overhauled after every flight costing ten's of millions of dollars. However, Henry Spencer a highly regarded expert on the history of space flight has said Rocketdyne studies show that with a lot of work to upgrade it, maintenance could be reduced to $750K per flight:

Spencer here said this would not be satisfactory for really large reductions in space costs. But this would be a reduction in SSME maintenance costs by 1 to 2 orders of magnitude, a major reduction in the costs for using the engine. The question is: how much would it cost to make the necessary upgrades to the engine?

Bob Clark

_________________Nanotechnology now can produce the space elevator and private orbital launchers. It now also makes possible the long desired 'flying cars'. This crowdfunding campaign is to prove it: